The present invention relates to a gas turbine operating at a pressure ratio of 20 or more. More particularly, this invention relates to a gas turbine capable of securely preventing drop of turbine efficiency due to shock wave loss.
A conventional gas turbine will be explained by referring to FIG. 7. Generally, a gas turbine comprises plural stages (four stages in this example) of stationary blades 1C to 4C arranged in a circle around a casing (blade ring or chamber) 1, and plural stages (four stages in this example) of moving blades 1S to 4S arranged in a circle around a rotor (hub or base) 2. FIG. 7 shows only the stationary blade 1C and moving blade 1S of the first stage, and the stationary blade 4C and moving blade 4S of the final stage, that is, the fourth stage (the area indicated by solid line ellipse).
Recently, in the mainstream of gas turbines, for example, the pressure ratio (expansion ratio) of all stages is large, that is, 20 or more (xcfx80xe2x89xa720), and the turbine inlet gas temperature is high, 1450xc2x0 C. or more (TITxe2x89xa71450xc2x0 C.). In such gas turbine, therefore, the pressure ratio of the final stage is 2.0 or more (xcfx80xe2x89xa72).
When the pressure ratio of the final stage is more than 2.0, as shown in FIG. 8, the average exit Mach number (M2) of the final stage moving blade 4S is in a range of 0.95 to 1.2 (0.95xe2x89xa6M2xe2x89xa61.2). Accordingly, by shock wave loss, the total pressure loss coefficient of the final stage moving blades 4S is in a suddenly increasing region. It means decline of turbine efficiency. In FIG. 8, the range of Mach number of M2 less than 0.95 in the arrow A direction shows the range of conventional gas turbine for power generation of low load or gas turbine for aircraft. The range of Mach number M2 greater than 1.2 in the arrow B direction shows the range of steam turbine.
Mechanism of decline of turbine efficiency due to shock wave is explained by referring to FIG. 9 and FIGS. 10(A) and (B). The final stage moving blade 4S is composed of a front edge 3, a rear edge 4, and a belly 5 and a back 6 linking the front edge 3 and rear edge 4. In FIG. 9, the blank arrow indicates the rotating direction of the final stage moving blade 4S.
When the pressure ratio of all stages is large, the pressure ratio of the final stage is also large, and the Mach number in the final stage moving blade 4S is large. In particular, as shown in the blade surface Mach number distribution in FIG. 10(B), the Mach number is large at the back 6 side of the final stage moving blade 4S. When the Mach number exceeds 1, as shown in FIG. 9, a shock wave 7 is generated. By generation of this shock wave 7, a boundary layer 8 (shaded area in FIG. 9 and FIG. 10(A)) is formed in a range from behind the shock wave 7 to the rear edge 4, at the back 6 side of the final stage moving blade 4S. This boundary layer 8 grows as the Mach number increases. By the growth of the boundary layer 8, the pressure loss increases, and the turbine efficiency is lowered. That is, by the boundary layer 8, the flow of the combustion gas G is disturbed, and the turbine efficiency is lowered.
The decline of turbine efficiency due to shock wave appears more prominently in the area of the tip side (for example, broken line oval area in FIG. 7), in particular, in the final stage moving blade 4S.
It is an object of this invention to provide a gas turbine capable of preventing securely decline of turbine efficiency due to shock wave loss in a gas turbine operating at a pressure ratio of 20 or more.
The gas turbine according to one aspect of this invention operates at a pressure ratio of 20 or more. Moreover, the gas turbine comprises a final stage of blades that includes a stationary blade and a moving blade. The moving blade is constructed such that the pressure difference at the downstream and upstream sides of said moving blade is 0.15 MPa or less.
The gas turbine according to another aspect of this invention operates at a pressure ratio of 20 or more. Moreover, the gas turbine comprises a final stage of blades that includes a stationary blade and a moving blade. In this gas turbine, a gauging ratio of said stationary blade is 0.9 or less. The gauging ratio is a ratio of a tip side gauging and a hub side gauging.
The gas turbine according to still another aspect of this invention operates at a pressure ratio of 20 or more. Moreover, the gas turbine comprises a final stage of blades that includes a stationary blade and a moving blade. In this gas turbine, an exit angle ratio of said stationary blade is 0.85 or more. The exit angle ratio is a ratio of a tip side exit angle and a hub side exit angle.
The gas turbine according to still another aspect of this invention operates at a pressure ratio of 20 or more. In this gas turbine, a duct wall in a portion of a specified distance from an end opposing said gas turbine, of a duct forming a diffuser passage communicating with a final exit side of said gas turbine is drawn parallel to or inside of a shaft of the gas turbine.
Other objects and features of this invention will become apparent from the following description with reference to the accompanying drawings.